The present invention generally relates an engine compressor and, more specifically, to a turbine engine compressor utilizing multiple component integration, thereby reducing the number of required engine components.
Referring to FIG. 1, there is shown a conventional small gas turbine engine 100. Intake air is taken into engine 100 as shown by arrows 102. The intake air passes through a compressor wheel 104. As air passes through the compressor section 106, it is accelerated outwards at high speeds. The accelerated air is slowed down in a diffuser 108, which comprises a ring of static vanes. A portion of the accelerated air may be used in a combustion chamber 110, and a portion of the accelerated air may be used to drive other cold turbines or to pressurize aircraft cabins. Engine 100 is powered by burning fuel in combustion chamber 110, heating the air flowing into engine 100, causing it to expand and gain kinetic energy. The hot gases generated by the combustion process drive one or more turbine wheels 112 to create mechanical power that may be used, for example, to drive compressor wheel 104.
Conventional compressor design includes multiple components, including an inner housing 114 to hold bearing requirements, an outer housing 116 to carry the carcass load, an inlet plenum 118, a separate bell mouth 120, a diffuser 108 and compressor wheel 104. A multiple component system makes it difficult to predict the structural behaviors due to thermal and mechanical loading during transient conditions. Holding tighter clearances between components becomes impossible due to the manufacturing tolerance build up among the various components. Such a multiple component approach will not meet the light weight requirements of high-performance aircraft engines, such as typical fighter jet engines.
The power and thermal management system (PTMS) of high-performance aircraft has technical challenges that require a novel approach in the design of the compressor module. The turbine engines of these high-performance aircraft operate as a typical auxiliary power unit (APU) on the ground in an open-loop, fired mode and transitions to a closed-loop mode using main propulsion engine bleed air for power. These two modes of operation both require optimal clearance control between the rotating group and the static structure of the engine.
As can be seen, there is a need for a new design concept for a compressor housing of a gas turbine engine that minimizes weight, cost and tolerance build up by utilizing multiple component integration to maintain the optimum engine performance under both the open-loop, fired mode and the closed-loop mode of operation.